JP2016125481A - Axial compressor rotor incorporating non-axisymmetric hub flowpath and splittered blades - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a compressor rotor that is operable with a sufficient stall range and an acceptable balance of aerodynamic and structural performance.SOLUTION: A compressor apparatus includes a rotor. The rotor includes: a disk mounted for rotation about a centerline axis, an outer periphery of the disk defining a flowpath surface having a non-axisymmetric surface profile; an array of airfoil-shaped axial-flow compressor blades extending radially outward from the flowpath surface, the compressor blades each having a root, a tip, a leading edge and a trailing edge; and an array of airfoil-shaped splitter blades alternating with the compressor blades, the splitter blades each having a root, a tip, a leading edge and a trailing edge. At least one of a chord dimension of the splitter blades at the roots thereof and a span dimension of the splitter blades is less than the corresponding dimension of the compressor blades.SELECTED DRAWING: Figure 1

Description

  The present invention relates generally to turbomachine compressors and, more particularly, to rotor blade stages of such compressors.

  The gas turbine engine includes a compressor, a combustor, and a turbine in series flow communication. The turbine is mechanically coupled to the compressor, and these three components define the turbomachine core. The core is operable in a known manner to generate hot pressurized combustion gases to operate the engine and to perform useful or mechanical work such as propulsion. One common type of compressor is an axial compressor with a plurality of rotors each containing a disk called a compressor blade having a row of axial airfoils.

  For reasons of thermodynamic cycle efficiency, it is generally desirable to incorporate a compressor with the highest pressure ratio (ie, inlet to outlet pressure ratio) that can be implemented. It is also desirable to include the least number of compressor stages. However, it is well known that there are interrelated aerodynamic limits for the maximum workable pressure ratio and mass flow rate through a given compressor stage.

  In order to reduce the mechanical stress in the disk, it is known to construct a disk having a non-axisymmetric “scalloped” surface profile. An aerodynamic disadvantageous side effect of this feature is that the rotor blade train passing through the flow zone is increased, the aerodynamic load level is increased, and air flow separation is facilitated.

  Accordingly, there remains a need for a compressor rotor that can operate with a sufficient stall range and an acceptable balance of aerodynamic and structural performance.

US Pat. No. 8,529,210

    According to one aspect of the present invention, a compressor apparatus having an axial flow rotor, the axial flow rotor being mounted to rotate about a central axis and having a non-axisymmetric surface profile on the outer periphery. Alternating with a disk defining a flow path surface, an array of airfoil axial flow compressor blades extending radially outward from the flow path surface, each having a root, a tip, a leading edge, and a trailing edge; An array of airfoil splitter blades having a root, a tip, a leading edge, and a trailing edge, wherein at least one of a chord dimension and a span dimension of the splitter blade is a correspondence of the compressor blade at the root of the splitter blade Smaller than the dimensions

  According to another aspect of the invention, the flow path surface includes concave scallops between adjacent compressor blades.

  According to another aspect of the invention, the scallop has a minimum radial depth adjacent to the root of the compressor blade and a maximum radial depth at a position approximately midway between adjacent compressor blades. Have.

  According to another aspect of the invention, each splitter blade is positioned approximately midway between two adjacent compressor blades.

  According to another aspect of the invention, the splitter blade is positioned such that the trailing edge of the splitter blade is at approximately the same axial position relative to the disk as the trailing edge of the compressor blade.

  According to another aspect of the invention, the span dimension of the splitter blade is 50% or less of the span dimension of the compressor blade.

  According to another aspect of the invention, the span dimension of the splitter blade is 30% or less of the span dimension of the compressor blade.

  According to another aspect of the present invention, the chord dimension at the root of the splitter blade is 50% or less of the chord dimension at the root of the compressor blade.

  According to one aspect of the present invention, a compressor apparatus having a plurality of axial flow stages, wherein at least one selected stage of the plurality of axial flow stages rotates about a central axis. Mounted on the disk and defining a flow path surface having a non-axisymmetric surface profile on the outer periphery and an airfoil axial flow compression extending radially outward from the flow path surface, each having a root, a tip, a leading edge, and a trailing edge An array of airfoil splitters and an array of airfoil splitter blades alternating with compressor blades, each having a root, a tip, a leading edge, and a trailing edge, the chord dimensions at the root of the splitter blade and the splitter blade At least one of the span dimensions is smaller than the corresponding dimension of the compressor blade.

  According to another aspect of the invention, the flow path surface includes concave scallops between adjacent compressor blades.

  According to another aspect of the invention, the scallop has a minimum radial depth adjacent to the root of the compressor blade and a maximum radial depth at a position approximately midway between adjacent compressor blades. Have.

  According to another aspect of the invention, each splitter blade is positioned approximately midway between two adjacent compressor blades.

  According to another aspect of the invention, the splitter blade is positioned such that the trailing edge of the splitter blade is at approximately the same axial position relative to the disk as the trailing edge of the compressor blade.

  According to another aspect of the invention, the span dimension of the splitter blade is 50% or less of the span dimension of the compressor blade.

  According to another aspect of the invention, the span dimension of the splitter blade is 30% or less of the span dimension of the compressor blade.

  According to another aspect of the present invention, the chord dimension at the root of the splitter blade is 50% or less of the chord dimension at the root of the compressor blade.

  According to another aspect of the invention, the selected stage is located in the rear half of the compressor.

  According to another aspect of the invention, the selected stage is the last stage of the compressor.

  The invention can best be understood by referring to the following description in conjunction with the accompanying drawings.

1 is a schematic cross-sectional view of a gas turbine engine incorporating a compressor rotor constructed in accordance with one aspect of the present invention. The perspective view of a part of rotor of a compressor apparatus. The top view of a part of rotor of a compressor apparatus. The rear elevation view of a part of the rotor of the compressor device. FIG. 5 is a side view taken along line 5-5 in FIG. FIG. 6 is a side view taken along line 6-6 of FIG.

  Referring to the drawings wherein like reference numerals represent like elements throughout the various views, FIG. 1 shows a gas turbine engine generally indicated at 10. The engine 10 has a longitudinal central axis 11 and, in the order of axial flow, a fan 12, a low pressure compressor or “booster” 14, a high pressure compressor (“HPC”) 16, a combustor 18, a high pressure turbine (“ HPT ”) 20 and low pressure turbine (“ LPT ”) 22. Overall, HPC 16, combustor 18 and HPT 20 define a core 24 of engine 10. HPT 20 and HPC 16 are connected to each other by an outer shaft 26. Overall, fan 12, booster 14 and LPT 22 define a low pressure system for engine 10. The fan 12, the booster 14 and the LPT 22 are connected to each other by an inner shaft 28.

  In operation, pressurized air from the HPC 16 is mixed with fuel in the combustor 18 and burned to generate combustion gases. Some work is extracted from these gases by the HPT 20 and drives the compressor 16 via the outer shaft 26. The remaining part of the combustion gas is discharged from the core 24 to the LPT 22. The LPT 22 takes work from the combustion gas and drives the fan 12 and booster 14 via the inner shaft 28. The fan 12 operates to generate a pressurized fan flow of air. The first part of the fan flow (“core flow”) flows into the booster 14 and the core 24, and the second part of the fan flow (“bypass flow”) exhausts through the bypass duct 30 surrounding the core 24. Is done. The illustrated embodiment is a high bypass turbofan engine, but the principles of the present invention are equally applicable to other types of engines such as low bypass turbofans, turbojets, and turboshafts.

  As used herein, the terms “axial” and “longitudinal” both refer to a direction parallel to the central axis 11, “radial” refers to a direction perpendicular to the axial direction, and It should be noted that “tangential direction” or “circumferential direction” refers to directions perpendicular to the axial direction and the radial direction. As used herein, the term “front” or “front” refers to a location that is relatively upstream of the air flow through or around the component, and the term “rear” Or “rear” refers to a location that is relatively downstream of the air flow through or around the component. The direction of this flow is indicated by the arrow “F” in FIG. These directional terms are used for convenience in the description only, and the specific orientation of the structure being described is not essential.

  The HPC 16 is configured for axial fluid flow, ie, fluid flow substantially parallel to the central axis 11. This is very different from a centrifugal or mixed flow compressor. The HPC 16 includes a plurality of stages, each of which includes a rotor having a row of airfoils or blades 32 (generally) mounted on a rotating disk 34 and a row of stationary airfoils or vanes 36. . The vane 36 serves to turn the air flow exiting the upstream row of blades 32 before entering the downstream row of blades 32.

  2-6 illustrate a portion of a rotor 38 constructed in accordance with the principles of the present invention and suitable for inclusion in the HPC 16. As an example, the rotor 38 can be incorporated into one or more of the rear half stages of the HPC 16, in particular the last or last stage.

  Rotor 38 includes a disk 40 having a web 42 and a rim 44. It will be appreciated that the completed disc 40 is an annular structure that is mounted for rotation about the central axis 11. The rim 44 has a front end 46 and a rear end 48. An annular channel surface 50 extends between the front end 46 and the rear end 48.

  A row of axial compressor blades 52 extends from the flow path surface 25. Each compressor blade includes a concave pressure side 58 that extends from a root 54 at the flow path surface 50 to a tip 56 and is joined to a convex suction side 60 at a leading edge 62 and a trailing edge 64. As best seen in FIG. 5, each compressor blade 52 connects a leading edge 62 and a trailing edge 64 with a span (or span dimension) “S1” defined as the radial distance from root 54 to tip 56. A chord (or chord dimension) “C1” defined as the length of the imaginary straight line. Depending on the specific design of the compressor blade 52, the chord C1 may be different at various locations along the span S1. In the present invention, the relevant reference dimension is the chord C1 at the root 54.

  As can be seen in FIG. 4, the flow path surface 50 is not a rotating body. Rather, the channel surface 50 has a non-axisymmetric surface profile. As an example of a non-axisymmetric surface profile, a concave bend or “scallop” 66 can be profiled between each adjacent pair of compressor blades 52. For comparison purposes, the dashed line in FIG. 4 shows a virtual cylindrical surface having a radius through the root 54 of the compressor blade 52. The curvature of the flow path surface has a maximum radius at the root 54 of the compressor blade 52 (i.e., the minimum radial depth of the scallop 66) and a minimum radius (i.e., approximately midway between adjacent compressor blades 52). It can be seen that the scallop 66 has a maximum radial depth “d”).

  During steady state or transient operation, this scalloped configuration is effective in reducing the magnitude of mechanical and thermal hoop stress concentrations at the airfoil hub intersection on the rim 44 along the flow path surface 50. . This contributes to the goal of achieving a long lifetime of acceptable components of the disk 40. An aerodynamic disadvantageous effect of scalloping the flow path 50 is an increase in the rotor passage flow area between adjacent compressor blades 52. This increase in the rotor flow area increases the aerodynamic load level, and as a result, on the suction side 60 of the compressor blade 52, in the inward portion near the root 54, and in the rear position (eg, leading edge). 62 to approximately 75% of chord C1) tends to cause undesirable flow separation.

  An array of splitter blades 152 extends from the flow path surface 50. One splitter blade 152 is disposed between each pair of compressor blades 52. In the circumferential direction, the splitter blade 152 is either intermediate between two adjacent compressor blades 52 or is circumferentially offset or is circumferentially aligned with the deepest portion d of the scallop 66. be able to. In other words, the compressor blades 52 and the splitter blades 152 are alternately arranged around the flow path surface 50. Each splitter blade 152 includes a concave pressure side 158 that extends from a root 154 to a tip 156 in the flow path surface 50 and is joined to a convex suction side 160 at a leading edge 162 and a trailing edge 164. As best seen in FIG. 6, each splitter blade 152 connects a span (or span dimension) “S2”, defined as the radial distance from root 154 to tip 156, with leading edge 162 and trailing edge 164. It has a chord (or chord dimension) “C2” defined as the length of the virtual straight line. Depending on the particular design of splitter blade 152, chord C2 may differ at various locations along span S2. In the present invention, the relevant reference dimension is the chord C2 at the root 154.

  The splitter blade 152 functions to locally increase the hub solidity of the rotor 38 and thereby prevent the aforementioned flow separation from the compressor blade 52. A similar effect is obtained by simply increasing the number of compressor blades 152 and thus reducing the spacing between the blades. However, this increases the aerodynamic surface area friction loss and has undesirable side effects that manifest as reduced aerodynamic efficiency and increased root weight. Accordingly, the size and location of splitter blade 152 can be selected to prevent flow separation while minimizing surface area. The splitter blades 152 are positioned such that their trailing edges 164 are at approximately the same axial position relative to the rim 44 as the trailing edge of the compressor blade 52. This can be seen in FIG. The span S2 and / or chord C2 of the splitter blade 152 may be a fraction of the corresponding span S1 and chord C1 of the compressor blade 52. These can be referred to as “partial span” and / or “partial chord” splitter blades. For example, span S2 can be equal to or less than span S1. Preferably, span S2 is about 50% or less of span S1 to reduce friction loss. More preferably, the span S2 is about 30% or less of the span S1 in order to achieve the minimum friction loss. As another example, chord C2 may be equal to or less than chord C1. More preferably, the chord C2 is about 50% or less of the chord C1 in order to achieve a minimum friction loss.

  The disk 40, compressor blade 52, and splitter blade 152 can be constructed of any material that can withstand the expected stress and environmental conditions during operation. Non-limiting examples of known suitable alloys include iron, nickel, and titanium alloys. 2-6, the disk 40, compressor blade 52, and splitter blade 152 are depicted as a generally unitary or monolithic structure. This type of structure can be referred to as a “bladed disk” or “blisk”. The principles of the present invention are equally applicable to rotors comprised of separate components (not shown).

  The rotor apparatus described herein with a splitter blade is a non-axisymmetric contoured hub that locally increases the solidity level of the rotor hub and locally reduces the aerodynamic load level of the hub. Suppresses the tendency of the rotor airfoil hub to try to flow separate in the presence of the channel surface. The use of partial span and / or partial chord splitter blades maintains the solidity levels of the middle and upper sections of the rotor that have not changed from their nominal values, thus maintaining the performance of the middle and upper airfoil sections. It is valid.

  The compressor rotor device has been described above. All features disclosed in this specification (including any appended claims, abstracts, and drawings), and all steps of any method or process so disclosed are thus Any combination can be combined except combinations where at least some of the features and / or steps are mutually exclusive.

  Each feature disclosed in this specification (including any appended claims, abstract, and drawings) serves the same, equivalent or similar purpose unless explicitly stated otherwise. Can be replaced with alternative features. Thus, unless expressly stated otherwise, each feature disclosed is one example only of a generic series of equivalent or similar features.

The present invention is not limited to the details of the one or more embodiments described above.
The invention includes any novel feature or any novel combination of features disclosed herein (including any appended claims, abstracts, and drawings), or any combination thereof, Can be extended to any new step or any new combination of any method or process steps disclosed.

F Flow direction C1 Chord S1 Span d Depth S2 Span C2 Chord 10 Engine 11 Axis 12 Fan 14 Booster 16 High pressure compressor 18 Combustor 20 High pressure turbine 22 Low pressure turbine 24 Core 26 Outer shaft 28 Inner shaft 30 Bypass duct 32 Blade 34 rotating disk 36 vane 38 rotor 40 disk 42 web 44 rim 46 front end 48 rear end 50 flow path surface 52 compressor blade 54 root 56 tip 58 pressure side 60 suction side 62 leading edge 64 trailing edge 66 scallop 152 splitter Blade 154 Root 156 Tip 158 Pressure side 160 Vacuum side 162 Front edge 164 Rear edge

Claims (20)

A compressor apparatus having an axial flow rotor (38), the axial flow rotor comprising:
A disk (34) mounted to rotate about a central axis (11) and defining a flow path surface (25, 50) having a non-axisymmetric surface profile on the outer periphery;
An array of airfoil axial flow compressor blades (52) extending radially outward from the channel surface and each having a root (54), a tip (56), a leading edge (62), and a trailing edge (64); ,
An array of airfoil splitter blades (152) alternating with the compressor blades, each having a root (154), a tip (156), a leading edge (162), and a trailing edge (164);
A compressor apparatus wherein at least one of a chord dimension at a root of the splitter blade and a span dimension of the splitter blade is smaller than a corresponding dimension of the compressor blade.   The apparatus of claim 1, wherein the flow path surface includes concave scallops between adjacent compressor blades.   2. The scallop according to claim 1, wherein the scallop has a minimum radial depth adjacent a root of the compressor blade and a maximum radial depth at a position approximately midway between adjacent compressor blades. Equipment.   The apparatus of claim 1, wherein each splitter blade is disposed approximately midway between two adjacent compressor blades.   The apparatus of claim 1, wherein the splitter blade is positioned such that a trailing edge of the splitter blade is substantially in the same axial position as the trailing edge of the compressor blade with respect to the disk.   The apparatus of claim 1, wherein a span dimension of the splitter blade is 50% or less of a span dimension of the compressor blade.   The apparatus of claim 1, wherein a span dimension of the splitter blade is 30% or less of a span dimension of the compressor blade.   The apparatus of claim 7, wherein a chord dimension at a root of the splitter blade is 50% or less of a chord dimension at a root of the compressor blade.   The apparatus of claim 1, wherein a chord dimension at a root of the splitter blade is 50% or less of a chord dimension at a root of the compressor blade. A compressor apparatus comprising a plurality of axial stages, wherein at least one selected stage of the plurality of axial stages is
A disk (34) mounted to rotate about a central axis (11) and defining a flow path surface (25, 50) having a non-axisymmetric surface profile on the outer periphery;
An array of airfoil axial flow compressor blades (52) extending radially outward from the channel surface and each having a root (54), a tip (56), a leading edge (62), and a trailing edge (64); ,
An array of airfoil splitter blades (152) alternating with the compressor blades, each having a root (154), a tip (156), a leading edge (162), and a trailing edge (164);
And at least one of a chord dimension at the root of the splitter blade and a span dimension of the splitter blade is smaller than a corresponding dimension of the compressor blade.   The apparatus of claim 10, wherein the flow path surface includes concave scallops between adjacent compressor blades.   11. The scallop according to claim 10, wherein the scallop has a minimum radial depth adjacent a root of the compressor blade and a maximum radial depth at a position approximately midway between adjacent compressor blades. Equipment.   The apparatus of claim 10, wherein each splitter blade is disposed approximately midway between two adjacent compressor blades.   The apparatus of claim 10, wherein the splitter blade is positioned such that a trailing edge of the splitter blade is at substantially the same axial position relative to the disk as a trailing edge of the compressor blade.   The apparatus of claim 10, wherein a span dimension of the splitter blade is 50% or less of a span dimension of the compressor blade.   The apparatus of claim 10, wherein a span dimension of the splitter blade is 30% or less of a span dimension of the compressor blade.   The apparatus of claim 16, wherein a chord dimension at a root of the splitter blade is 50% or less of a chord dimension at a root of the compressor blade.   The apparatus of claim 10, wherein a chord dimension at a root of the splitter blade is 50% or less of a chord dimension at a root of the compressor blade.   The apparatus of claim 10, wherein the selected stage is disposed in a rear half of the compressor. The apparatus of claim 10, wherein the selected stage is the last stage of the compressor.